1. Field of the Invention
The present invention relates to apparatuses and methods for the implementation of electric propulsion in spacecraft. Particularly, the present invention relates to mounting solutions for Hall effect thrusters to mitigate harmful thruster plume effects.
2. Description of the Related Art
In recent years, electric propulsion systems, e.g. xenon ion propulsion systems, have been implemented in spacecraft. Current high power applications of this thruster now make it possible to accomplish both orbit raising and station keeping. This combination may yield an all electric spacecraft with a propulsion system several times lighter than a chemical system. However, electric thruster plumes are typically wide and avoiding spacecraft structure can be problematic. Particularly, Hall effect electric thrusters (operating without front grids) are known to have widely dispersed plumes resulting in a significant potential for negative plume interactions.
Conventionally, electric propulsion thrusters, including Hall effect thrusters, are mounted either on the spacecraft body or to other primary structure. Within this general class there are offset orthogonal mounts, e.g., thruster pairs that fire parallel to each other through lines of action at equal offsets from the center of gravity. See, e.g. Director, Operational Test & Evaluation FY2001 Annual Report, Advanced Extremely High Frequency (AEHF) Satellite Communications System, which is incorporated by reference herein.
On the other hand, canted mounts can be designed such that both thrusters fire directly through the center of gravity at an angle. The canted angle allows the plumes to be directed away from each other and spacecraft structure, however there is loss for each thruster based on the cosine of one half the angle between the intersecting thruster lines. See, e.g. G. Saccoccia et al., Electric Propulsion: A Key Technology for Space Missions in the New Millennium, ESA Bulletin, February 2000, which is incorporated by reference herein.
In addition, hybrid mounting designs can provide canted offsets (canted thrusters which are offset and do not fire through the center of gravity) and obtain some of the benefits each fundamental type. That is, such designs can have plumes directed away from each other and spacecraft structure by the canted angles and further displaced from each other and spacecraft structure by the offsets, however, there will be a loss based upon the severity of the canted angle.
With any of these conventional arrangements, however, plumes from electric propulsion thrusters can have adverse effects on antenna performance, loss of radiator panel efficiency and loss of solar cell efficiency that are directly related to the location of electric thrusters on spacecraft. These effects can occur through direct sputtered material deposition as well as secondary back flow charge exchange plasma that focuses ion impingement on spacecraft surfaces.
Generally, the location of such thrusters is driven by the control efficiency gained from their firing orientation being through the spacecraft center of gravity. The proximity of these thrusters to the body of the spacecraft, as the result of this orientation requirement, ensures that there will be plume interaction that will increase the risk of performance degradation over the typical satellite mission life. This effect is of particular concern when considering a typical 10 to 15 year communication satellite life and the lack of on-orbit data that can correlate to the known effects observed in ground tests.
In view of the foregoing, there is a need for spacecraft designs which mitigate the adverse effects of electric propulsion thruster plumes. Further, there is a need for such designs to be made compatible with existing subsystems, minimizing additional separate structure and mechanisms. As discussed hereafter, the present invention meets these and other needs.